This invention relates to composite materials and more particularly to a protective coating for composite materials.
Graphite-epoxy materials and other composite materials containing graphite fibers, are widely used in spacecraft to form structural members and other parts, because of their unique combination of very high mechanical strength, light weight (low physical density) and good resistance to elevated temperatures. These materials display three disadvantages, however, in such applications. A first disadvantage is due to their dark coloration, which is generally gray to black; as a result, these materials tend to be strong absorbers of solar radiation, which leads to difficulties in thermal stabilization of the spacecraft made of this material.
A second disadvantage is due to their organic material content. They are susceptible to both physical degradation (erosion damage) and chemical degradation (carbonization and other compositional changes) caused by bombardment by ozone, ionic and atomic oxygen, and other gaseous or high energy particles under orbital conditions, especially at lower altitudes. Oxygen presents the worst problem because of its high chemical affinity for most organic materials, including carbon itself. These composite materials are also easily burned through by impinging laser beams, since they are not truly refractory. A third disadvantage is that, under ordinary conditions of temperature, and humidity, these materials, being somewhat porous, tend to absorb comparatively large quantities of water (up to 1.5% by weight) after part fabrication and installation aboard a spacecraft. The water absorbed on the surface spreads quickly throughout the epoxy matrix, by diffusion processes enhanced by capillary effects along the embedded graphite fibers. In effect, this water simulates the water of hydration in an inorganic crystalline material such as copper sulfate. Subsequently, in the vacuum of space, much of the moisture is gradually desorbed in the form of water vapor. This exudation of water can cause spacecraft contamination and, much worse, structural damage in the form of component weakening, shrinkage, and warping, with resultant distortion of the spacecraft installations. Similarly, if the water freezes within the epoxy matrix, due to the low temperatures, it can cause cracking or fracturing of the component. For example, such disturbances would constitute a serious drawback in optical systems which require precision alignment for proper performance.
It is known that thermal stabilization can be achieved by depositing an opaque coating of a highly reflective material such as silver, aluminum, magnesium or gold on the surface. While this may correct the heating problem, the chemical degradation and physical damage caused by bombardment with zone, ionic and atomic oxygen and other gaseous and high energy particles under low earth orbit conditions, can cause erosion and destruction of such a coating. It has been proposed to provide a metal oxide protective coating over the reflective coating and also a metal oxide subcoating to smooth out the surface for the reflective material coating. This thin coating may be of a very chemically stable, optically transparent, hard refractory metal oxide such as Al.sub.2 O.sub.3 (aluminum oxide), SiO.sub.2 (silicon dioxide), TiO.sub.2 (titanium dioxide), or MgO (magnesium oxide). These oxides are very resistant to mechanical damage (scratching and pitting) because of their high Brinell hardness.
These configurations provide protection against solar radiation and the physical degradation and chemical degradation caused by bombardment by ozone, ionic and atomic oxygen, and other gaseous or high energy particles under low earth orbit conditions. These configurations do not provide for protection from earth-based absorption of large quantities of moisture which is subsequently gradually desorbed in the vacuum of space in the form of water vapor. This can cause spacecraft contamination as well as structural damage or misalignment.